Pratt-whitney Canada Corp

Raytheon Technologies Corporation

Daniel Lecuyer

Panagiota Tsifourdaris

Robert Huszar

Michael Papple

Othmane Leghzaouni

Niloofar Moradi

Abdulhalim Twahir

Remo Marini

Franco Di Paola

Michael Leslie Clyde Papple

Francis Gemme

Guy Lefebvre

Timothy J Jennings

Ghislain Plante

Mohammed Ennacer

David A Niezelski

Domenico Di Florio

Nicolas Grivas

Rene Paquet

Bruno Chatelois

Alain Bouthillier

Masoud Roshan Fekr

John R Battye

Jasrobin Grewal

Alain Carrieres

Alan Juneau

Michael Paolucci

Stephane Bigras

Patricia Phutthavong

Patrick Peloquin

Jonathan Dorai

Farzad Ashrafi

Francois Caron

Alan Bouthillier

Andrew Shapiro

XingYun Haggard

Jocelyn Menard

A blade for mounting on a rotor disc of a turbine engine includes platform segments between a airfoil and blade root and extending laterally from the blade, the platform segment having a curved profile defining a platform recess on a root side of the platform segment, the recess configured to cooperate with recesses of adjacent blades to define a blade pocket. A recessed region is defined in a rear surface of the blade root, and is configured to receive a complementary projection of the rotor disc upon mounting the blade on the rotor disc. The recessed region has a surface configured to abut a corresponding surface on the projection of the rotor disc to form a face seal. Methods of manufacturing such a rotor disc and of sealing a circumferential joint between such blades and the rotor disc are also disclosed.

Blade for a rotor assembly of a turbine engine

An airfoil is provided that includes a first end, a second end, a leading edge, a trailing edge, a pressure side and a suction side. The leading edge and the trailing edge are joined by the pressure side and the suction side to provide an exterior airfoil surface extending in a spanwise direction from the first end of the airfoil to the second end of the airfoil. The exterior airfoil surface are formed in conformance with a plurality of cross-section profiles of the airfoil described by a set of Cartesian coordinates set forth in Table 1. The Cartesian coordinates are provided by an axial coordinate scaled by a local axial chord, a circumferential coordinate scaled by the local axial chord, and a span location. The local axial chord corresponds to a width of the airfoil between the leading edge and the trailing edge at the span location.

Turbine engine airfoil

A fir tree structure is provided and includes a part extending radially inwardly from a rotating part and having an inward taper. The part has a spline profile forming fixing lobes and necks, which are interleaved with the fixing lobes. The spline profile has constantly changing radii of curvature characterized in that the radii of curvature are relatively reduced in high stress locations and relatively increased in low stress locations.

Turbine rotor dovetail structure with splines

A blade for mounting on a rotor disc of a gas turbine engine includes platform segments between a airfoil and blade root and extending laterally from the blade, the platform segment having a curved profile defining a platform recess on a root side of the platform segment, the recess configured to cooperate with recesses of adjacent blades to define a blade pocket. A recessed region is defined in a rear surface of the blade root, and is configured to receive a complementary projection of the rotor disc upon mounting the blade on the rotor disc. The recessed region has a surface configured to abut a corresponding surface on the projection of the rotor disc to form a face seal. Methods of manufacturing such a rotor disc and of sealing a circumferential joint between such blades and the rotor disc are also disclosed.

Method of manufacturing a rotor disc for a turbine engine

A turbine blade for a gas turbine engine has an airfoil including leading and trailing edges joined by spaced-apart pressure and suction sides to provide an external airfoil surface extending from a platform in a spanwise direction to a tip. The external airfoil surface is formed in substantial conformance with multiple cross-sectional profiles of the airfoil defined by a set of Cartesian coordinates set forth in Table 1, the Cartesian coordinates provided by an axial coordinate scaled by a local axial chord, a circumferential coordinate scaled by a local axial chord, and a span location.

Turbine blade airfoil profile

An airfoil for a gas turbine engine, which has an airfoil body extending in a spanwise direction and in a chordwise direction, a platform located at an inner end and/or an outer end of the airfoil body, and a fillet at a junction between the airfoil body and the platform. The fillet has a radius distribution at a given chordwise location, the radius distribution varying from the platform to the airfoil body in the spanwise direction. The radius distribution defines a local minimum, the radius of the fillet at the given chordwise location increasing from the local minimum along the spanwise direction toward both of the airfoil and the platform. A local maximum of the radius distribution is offset from the local minimum along the spanwise direction, the radius decreasing from the local maximum along the spanwise direction toward both of the airfoil and the platform.

Airfoil having a spline fillet

A turbine vane for a gas turbine engine has an airfoil including leading and trailing edges joined by spaced-apart pressure and suction sides to provide an external airfoil surface. The surface is formed in substantial conformance with multiple cross-sectional profiles of the airfoil defined by a set of Cartesian coordinates set forth in Table 1, the Cartesian coordinates provided by an axial coordinate scaled by a local axial chord, a circumferential coordinate scaled by a local axial chord, and a span location.

Turbine vane airfoil profile

A turbine blade for a gas turbine engine has: an airfoil extending along a span from a base to a tip and along a chord from a leading edge to a trailing edge, the airfoil having a pressure side and a suction side, a tip pocket at the tip of the airfoil, the tip pocket at least partially surrounded by a peripheral tip wall defining a portion of the pressure and suction sides; at least one internal cooling passage in the airfoil and having at least one outlet communicating with the tip pocket; and a reinforcing bump located on the pressure side of the airfoil and protruding from a baseline surface of the peripheral tip wall to a bump end located into the tip pocket, the reinforcing bump overlapping a location where a curvature of a concave portion of the pressure side of the airfoil is maximal.

Turbine blade for a gas turbine engine

A turbomachine airfoil element comprises an airfoil having: an inboard end; an outboard end; a leading edge; a trailing edge; a pressure side; and a suction side. A span between the inboard end and the outboard end is 1.75-2.20 inches. A chord length at 50% span is 1.05-1.35 inches. At least two of: a first mode resonance frequency is 2400&#xb1;10% Hz; a second mode resonance frequency is 4950&#xb1;10% Hz; a third mode resonance frequency is 7800&#xb1;10% Hz; a fourth mode resonance frequency is 8700&#xb1;10% Hz; and a fifth mode resonance frequency is 12500&#xb1;10% Hz.

Turbomachine blade

The gas turbine engine rotor assembly includes a rotor disc with a plurality of sealing tabs projecting radially out from a peripheral surface of the rotor disc in the rear end portion thereof. A said sealing tab is disposed at a tip portion of a fixing member of the disc, formed between pairs of blade root slots, adjacent a trailing edge of the rotor disc. The sealing tabs help to reduce the leakage of secondary air out the back of a blade pocket defined between adjacent blades mounted on the rotor disc.

Rotor assembly with blade sealing tab

An annular shroud assembly for a gas turbine engine is provided. The annular shroud assembly includes a plurality of first shroud segments having a same first arcuate length within a tolerance, at least one second shroud segment having a second arcuate length different than the first arcuate length and outside the tolerance, a plurality of first intersegment gaps between adjacent first segments, the first intersegment gaps having a circumferential dimension within a desired controlled range of dimensions, and at least two second intersegment gaps between opposed ends of the second segment and adjacent first segment. The first and second intersegment gaps are within the desired controlled range. Methods for controlling intersegment gaps between a plurality of shroud segments forming an annular shroud assembly and for forming such annular shroud assembly are also provided.

Shroud segment assembly intersegment end gaps control

A partially coated blade for a gas turbine engine, including a fillet surface surrounding the airfoil section and connecting it to the platform section. A radially outermost portion of the pressure side and leading edge is covered by a thermal barrier coating. This portion extends radially from a first limit to the blade tip. The first limit is located at a radial distance from the platform of at most 21% of the maximum span. The fillet surface is free or substantially free of the thermal barrier coating. In another embodiment, a second portion of the pressure side and of the leading edge is free or substantially free of the thermal barrier coating, extending radially from the platform section to a second limit located a radial distance from the platform section corresponding to at least 5% of the maximum span. A method of applying a thermal barrier coating is also discussed.

Partially coated blade

A turbine blade for a gas turbine engine comprises an airfoil having a pressure side, a suction side, a span direction and a chord-wise direction. The airfoil has an airfoil span on a pressure line being a projection of the stacking line onto the pressure side. The airfoil has a plurality of chords extending between a leading edge and a trailing edge of the airfoil. A generally round dimple is disposed on the pressure side. The dimple is contained in an area extending on the stacking line between 0% and 23% of the airfoil span from the inner end, and in the chord-wise direction between 0% of a first chord and 82% of a second chord from the leading edge. The dimple is configured to initiate fracture of the blade at a predetermined speed of rotation. A method of preventing rupture of a disk of a turbine rotor is also presented.

Growing community of inventors

St-Bruno-de-Montarville, Canada

Daniel Lecuyer